1. Field of the Invention
This invention relates to autonomous unmanned space flight systems and planetary landers, and more specifically to a planetary lander for executing a discrete landing sequence that can land softly with pinpoint accuracy, detect and avoid hazards, and relocate by taking off and flying to different sites. Hazard avoidance and relocation enhance the mission capability of the lander.
2. Description of the Related Art
The mission of the National Aeronautics and Space Agency (NASA) is to pioneer the future in space exploration, scientific discovery and aeronautics research. Although manned exploration in the form of the moon landings, the space shuttle missions, international space station and possible mission to Mars garner most of the attention, unmanned exploration is critical to NASA's mission. Unmanned exploration is less risky, more cost effective and can perform missions beyond the reach of manned exploration. Unmanned planetary landers such as Viking, the Mars Polar Lander, Phoenix, Survey or etc. provide the capability to land softly on a planet, asteroid or other body in the solar system (hereinafter “planet”) and perform important planetary science.
Conventional all-propulsive autonomous planetary landers ignite liquid fueled thrusters at the terminal landing altitude after aeroshell and parachute separation (in atmosphere, aka Mars) or solid rocket motor (SRM) separation (airless bodies, aka the Moon) and burn continuously until the last meter before landing to execute a continuous landing sequence. Some landers forgo the SRM burn and rely solely on liquid thrusters. The SRM/parachute and liquid-fuel thrusters are designed to gradually decelerate the lander starting at a relatively high altitude (few kms) so that the lander falls slowly through the near surface constant gravity field and lands almost directly beneath the point at which the thrusters are ignited. During most of the descent, the thrusters have been specifically designed to produce a thrust that is a fraction of the specific gravity of the planet so that the lander resists the pull of gravity and falls very gently. At the last moment, when most of the fuel has been depleted the thrust briefly exceeds the planet's gravity allowing the lander to stop briefly before free falling the last few meters for a soft landing on the surface. The liquid propulsion system uses mono or bi-propellant thrusters, usually in a cluster configuration, that produce just enough thrust for the controlled descent over extended burn times, e.g. 90 seconds. For example, the Phoenix Mars lander uses a twelve thruster cluster to produce a total thrust-to-mass ratio (T/M) (thrust to total wet lander mass) in m/sec2 of about 3:1 during initial free fall and increasing to about 15:1 (when most of the fuel has been depleted) momentarily to stop. The configuration of the propulsion system will depend on the planet's actual gravity, but the principles are the same. Mars landers will require more fuel and thrusters to produce greater thrust due to Mars gravity (as compared to lunar) for a similar continuous burn time. In addition, during descent the thrusters are usually off-pulse modulated to balance the center of mass and reduce the time averaged thrust to maintain the desired T/M ratio as the propellant is depleted.
Essentially all of the lander's change in velocity (ΔV) capability (total integrated thrust from the SRM and liquid to remove lander velocity) is used to land softly on the planet's surface. The lander's fuel mass fraction (FMF) (ratio of fuel mass to total lander mass) is high, approximately 40% or greater. The remaining available mass is allocated to payload (e.g. scientific instruments) and required lander systems and even than the payload mass fraction (PFM) (ratio of payload mass to total dry mass) is only about 5-10% typically. Landers may remove a portion of the known guidance error caused by SRM burn or the parachute but do not address the unknown navigation errors. The lander does not have the fuel margin, navigation measurement capability or T/M ratio to remove unknown navigation errors. The additional fuel alone would increase the FMF to the point that there would be no remaining mass left to allocate to payload thereby entirely defeating the purpose of the mission. Even if fuel were available, the limited T/M capability is inadequate to efficiently remove navigation error. Conventional landers are designed to land the payload on the surface of the planet with minimum risk in areas know a priori (via remote sensing images) to be free of hazards such as rocks.
Typical state-of-the art landers have an error of approximately a 1 km×3 km ellipse for a lunar landing and approximately 30 km×230 km ellipse for a landing on Mars. The early SRM burn to remove lander velocity (no atmosphere) and the use of the aeroshell to penetrate the atmosphere and the parachute to slow the lander (atmosphere) induces considerable error that cannot be removed by the lander due to the lack of ΔV capability. This lack of precision delivery capability limits landing sites to flat, relatively uninteresting areas from a planetary science perspective in order to provide a low-hazard landing site. As mapping improves (higher resolution images) hazards are discovered and it becomes more difficult to find suitable landing sites that satisfy the risk averse mission planners. The instrument package may be incorporated in a rover that can drive to more desirable sites but this increases risk (rover survives landing and drives for many days to the site), mass and cost.